Many supersonic aircraft employ gas turbine engines that are capable of propelling the aircraft at supersonic speeds. These gas turbine engines, however, generally operate on subsonic flow in a range of about Mach 0.3 to 0.6 at the upstream face of the engine. In supersonic applications, a nacelle is used to encompass the engine and incorporates an inlet and a nozzle. The inlet decelerates the incoming airflow to a speed compatible with the requirements of the gas turbine engine. To accomplish this, a supersonic inlet is comprised of a compression surface and corresponding flow path, used to decelerate the supersonic flow into a strong terminal shock. Downstream of the terminal shock, subsonic flow is further decelerated using a subsonic diffuser to a speed corresponding with the in-flow requirements of the gas turbine engine. The exhaust from the engine is then accelerated again using the nozzle.
Traditional supersonic propulsion system design methods minimize the diameter and structural weight of the nacelle while maximizing gross thrust. In doing so, the amount of flow captured by the inlet is limited to only that demanded by the engine with an additional small amount for nacelle purge and cooling. A measurement of inlet operation efficiency is the total pressure lost in the air stream between the entrance side and the discharge side of the inlet. The total pressure recovery of an inlet is defined by a ratio of the total pressure at the discharge to the total pressure at the free stream. Maximizing the total pressure recovery leads to maximizing gross engine thrust, thus improving the performance of the propulsion system.
FIG. 1 schematically illustrates a cross-sectional view of a traditional nacelle 10, having an external compression inlet 11 and nozzle 14 surrounding an engine 16. The external compression inlet 11 compresses and decelerates the supersonic flow to the face of the engine 16. The inlet 11 includes the leading edge 12 of the compression surface and the cowl 13 forming the inlet opening of the inlet 11. The output of the engine 16 is then accelerated by the nozzle 14, creating the necessary thrust to propel the aircraft at supersonic speeds. The nacelle 10 is often designed to cover around the protruding engine parts 18, which may include engine components such as gear boxes and other components known to those of skill in the art. As shown in FIG. 1, the engine 16 may be a conventional turbofan type engine featuring approximately 15,000 lbf of maximum takeoff thrust and a moderate fan-to-compressor flow ratio of 3.
Unfortunately, the traditional nacelle design for a supersonic engine configuration often generates strong shocks off the supersonic inlet and from the body of the nacelle. A traditional approach to supersonic inlet design typically employs shock-on-lip focusing. As would be understood by those of skill in the art, shock-on-lip focusing involves designing a compression surface configuration of an external compression inlet such that the inlet-generated shocks (that occur at a supersonic design cruise speed) meet at a location immediately forward of the cowl highlight or the cowl lip.
FIG. 2 schematically illustrates the bottom half of the cross-sectional view in FIG. 1 and how shock waves and expansion regions are generated by the nacelle 10 and how the shape of the nacelle 10 may generate additional shock waves and expansion regions. As is understood by those of skill in the art, these shock waves can coalesce into a stronger shock wave as the shock waves propagate away from the aircraft during supersonic flight. These shock waves can also propagate into the aircraft surface, creating localized regions of interference drag. The shock wave 20 is generated by the leading edge 12 of the initial compression surface of the inlet 11. The wave 20 may coalesce with the shock wave 22, generated by the cowl 13, and potentially the shock waves 24 and 26. These shock waves may then coalesce with shock waves propagating from the airframe itself, eventually creating a sonic boom as heard at ground level.
The shock wave 22 is often referred to as the cowl shock, the strength of which may be directly related to the cowling angle A. In addition, any increase in cowling angle results in additional inlet frontal area, which increases inlet drag as speed increases. This adverse trend is a key reason why conventional external compression inlets lose viability at high supersonic Mach numbers. Other shock waves, such as shock wave 24, and expansion regions, identified in region 25, are often caused by changes in the shape and diameter of the nacelle 10, especially as the nacelle attempts to cover the protruding engine parts 18. The shock wave 26 is generated off the trailing edge 15 of the nozzle. As is understood by those of skill in the art, the strength of this shock wave 26 is proportional to the nozzle cowling angle B, often referred to as the nozzle boat tail angle.
Unfortunately, these shock waves combine with those from the airframe to create a louder overall sonic boom signature and more interference drag between the nacelle and the remainder of the vehicle. The stronger the shock waves, the more difficult they become to control and attenuate and the more likely they are to produce additional drag and sonic boom noise.
One way to control drag, as discussed in U.S. Pat. No. 6,793,175 to Sanders, involves configuring the inlet to minimize the shape and size of the cowl. The configuration of the inlet initially resembles a circumferential sector of an axisymmetric intake, but switches the location of compression surface to the outer radius and disposes the cowling on the inner radius in a higher performance, 3-D geometry. The fact that the cowl is located on the inner radius reduces the physical arc of the cowl. Problems with this method include the aircraft integration challenges created by the 3-D geometry, such as the fact that the cross-sectional shape may be more difficult to integrate from a packaging perspective compared to an equivalent axisymmetric design for podded propulsion systems. In addition, the complex inlet shape is likely to create complex distortion patterns that require either large scale mitigating techniques in the subsonic diffuser or the use of engines with more robust operability characteristics.
Another way to control drag by reducing the cowl lip angle is based on decreasing the flow turn angle by increasing the inlet terminal shock Mach number. The improvement in drag reduction is often negated by the reduction in pressure recovery resulting from the stronger terminal shock. In addition, increasing the terminal shock Mach number at the base of the shock also encounters significant limitations in practice due to viscous flow effects. Higher terminal shock Mach numbers at the base of the shock aggravate the shock-boundary layer interaction and reduce shock base boundary layer health. The increase in shock strength in the base region also reduces inlet buzz margin, reducing subcritical flow throttling capability. Additionally, the increase in terminal shock Mach number will most likely require complex boundary layer management or a complex inlet control system.
Inlet compression surfaces are typically grouped into two types: straight or isentropic. A straight surface has a flat ramp or conic sections that produce discrete oblique or conic shocks, while an isentropic surface has a continuously curved surface that produces a continuum of infinitesimally weak shocklets during the compression process. Theoretically, a traditional isentropic compression surface can have better pressure recovery than a straight surface designed to the same operating conditions, but real viscous effects can reduce the overall performance of the isentropic surface inlets and result in poorer boundary layer health.
FIG. 3 schematically represents a perspective view of a engine arrangement 30 representative of a high specific thrust military turbofan engine of approximately 11,000 lbf maximum takeoff thrust class (non-afterburning). The arrangement 30 may include a nacelle 32, having a traditional inlet 34 and nozzle 36. As can be seen from FIG. 3, the nacelle must be configured to encompass the protruding parts 40 of the engine 38. Moreover, the non-optimal matching between the intake area and the maximum nacelle diameter creates a large forward cowling profile that results in high drag and strong shock generation. Likewise, the non-optimal matching between exhaust area and maximum nacelle diameter causes a large nozzle boat tail angle, resulting in high drag and strong expansion and re-shock.
FIG. 4 schematically illustrates a perspective view of the engine 38 from FIG. 3 installed on the vertical stabilizer of a supersonic aircraft 42. The nacelle 32 is configured to encompass the protruding engine parts 40, creating a generally asymmetric configuration, which as discussed above may contribute to the generation of shock waves and the strength of a resulting sonic boom. While such performance may be acceptable for military aircraft or other such applications, the generation of strong sonic booms in the civil aviation arena is undesirable.